CN112363524B - Reentry aircraft attitude control method based on adaptive gain disturbance compensation - Google Patents
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
- G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
- G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
- G05D1/10—Simultaneous control of position or course in three dimensions
- G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft
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Abstract
A reentry aircraft attitude control method based on adaptive gain disturbance compensation comprises the following steps: establishing a reentry flight kinematics and dynamics model of the reentry aircraft facing control; step two: establishing a fixed time convergence disturbance compensation observer, and observing the state and disturbance item of the reentry vehicle; step three: and designing an integral sliding mode controller, introducing an observer disturbance observation item into the integral sliding mode controller, and adopting a double-layer self-adaptive gain strategy to carry out self-adaptive adjustment on control gain. The novel fixed time convergence disturbance compensation observer designed by the invention can ensure that the observation error is rapidly converged to zero in the fixed time, has good noise suppression capability, and greatly improves the anti-interference capability of the aircraft.
Description
Technical Field
The invention relates to a reentry vehicle attitude control method based on adaptive gain disturbance compensation, and belongs to the technical field of guidance and control.
Background
With the development of modern technology, the military value and civil value of reentry vehicles are increasingly prominent. The reentry vehicle has various execution tasks and complex flight environment, so that the reentry gesture control problem has the characteristics of multiple variables, quick time variation, strong coupling, strong nonlinearity, uncertainty of parameters and the like, and the design difficulty of a control system is greatly improved. In recent years, home and abroad scholars have developed intensive researches on reentry gesture control problems, and common control system design methods mainly comprise self-adaptive control, active disturbance rejection control, robust control, sliding mode control and the like, but currently, the rapid stability performance, gesture tracking precision, disturbance rejection performance and the like of a gesture control system are still required to be further researched and explored for a plurality of key technologies and complex new problems.
Disclosure of Invention
The technical solution of the invention is as follows: aiming at the problem of controlling the attitude of the reentry vehicle in a complex flight environment, the model uncertainty and the influence of various external disturbances are fully considered, and the reentry vehicle attitude control method based on adaptive gain disturbance compensation is provided, so that the attitude control precision and the disturbance rejection capability are improved.
The technical solution of the invention is as follows: the reentry vehicle attitude control method based on the adaptive gain disturbance compensation comprises the following three steps:
step one: fully considering factors such as model unmodeled dynamics, model uncertainty, pneumatic parameter uncertainty, moment disturbance and the like, establishing a reentry flight kinematics and dynamics model of the reentry aircraft facing control, ensuring that reentry flight characteristics can be truly reflected, and laying a foundation for control system design;
step two: establishing a novel fixed time convergence disturbance compensation observer, observing the state and disturbance item of the reentry vehicle, and proving the stability and convergence of the observer by adopting Lyapunov stability theorem;
step three: the integral sliding mode controller is designed, an observer disturbance observation item is introduced into the integral sliding mode controller, meanwhile, a double-layer self-adaptive gain strategy is adopted to carry out self-adaptive adjustment on control gain, and the stability and convergence of the designed controller are proved by adopting a Lyapunov stability theorem.
Further, the reentry vehicle is control-oriented reentry kinematic and dynamic model, specifically:
wherein θ= [ α, β, μ ]] T α, β, μ represent angle of attack, sideslip angle, and roll angle, respectively; omega= [ omega ] x ,ω y ,ω z ] T ，Is the first derivative; omega x ,ω y ,ω z Representing three-channel attitude angular velocity; b represents a control matrix, u represents a control input vector; d represents disturbance moment, and ΔJ represents a moment of inertia error matrix; the matrix R, J, Ω expressions are:
wherein J is x ，J y ，J z Three-channel main moment of inertia, J xy ，J xz ，J yz Respectively the product of inertia between the two channels.
Further, the reentry flight kinematics and dynamics model is further simplified to obtain:
wherein the method comprises the steps ofRepresenting the total disturbance term, wherein the boundary of the total disturbance term is an unknown constant, namely the I delta I is less than or equal to D.
Further, the fixed time convergence disturbance compensation observer is of the form:
wherein z is 1 ,z 2 The observed amounts of ω, Δ,z respectively 1 ,z 2 Is the derivative of the power exponent alpha i And beta i The construction mode of (a) is as follows: alpha is E (1-epsilon, 1) and alpha i ＝iα-(i-1)，i＝1,2，β∈(1,1+ε 1 ) And beta is i ＝iβ-(i-1)，ε>0 and epsilon 1 >0 are all positive small amounts; gain kappa i And k i The matrix is constructed to satisfy the Helviz condition, where
Further, let the observation error be e 1 ＝z 1 -ω，e 2 ＝z 2 - Δ, then the error equation of the fixed time converging disturbance compensation observer is as follows:
here, theE is 1 ,e 2 Is a derivative of (a).
Further, the third step: an integral sliding mode controller is designed, an observer disturbance observation item is introduced into the integral sliding mode controller, and meanwhile, a double-layer self-adaptive gain strategy is adopted to carry out self-adaptive adjustment on control gain, specifically:
(1) For reentry flight kinematics and dynamics models, virtual control quantity omega is designed c Is that
ω c ＝R -1 (θ c -k 1 e θ )
Where the gain k 1 For the parameters to be designed e θ ＝θ-θ c For tracking error, θ c Representing a desired attitude angle command;
(2) The integral sliding mode controller u is designed for reentry flight kinematics and dynamics models as follows:
u＝u 1 +u 2
wherein u is 1 For canceling known system constitution and external disturbances, u 2 For compensating the estimation error;
u 1 in the form of，B + A pseudo-inverse matrix representing the control matrix B;
u 2 in the form of
Wherein u is 2n ＝-ρ n B + Je ω To compensate for nominal part of the error, e ω Tracking error e for angular velocity ω ＝ω-ω c ，ρ n As a variable to be designed,for integrating the sliding surface, ρ (t) is the tableThe outer layer adaptive control gain is shown, and r (t) represents the inner layer adaptive control gain in the form:
wherein sign (·) is a sign function, r 0 Is the initial value of r (t), and gamma is the parameter to be designed and satisfies gamma>0, and delta (t) has the form:
wherein the method comprises the steps ofIs the filter output of equivalent control, alpha and epsilon are parameters to be designed and satisfy 0<α<1,ε>0。
Furthermore, the invention also provides a reentry aircraft attitude control system, which comprises:
a dynamics model modeling module: establishing a reentry flight kinematics and dynamics model of the reentry aircraft facing control;
the observer building module: establishing a fixed time convergence disturbance compensation observer, and observing the state and disturbance item of the reentry vehicle;
the controller design module: and designing an integral sliding mode controller, introducing an observer disturbance observation item into the integral sliding mode controller, and adopting a double-layer self-adaptive gain strategy to carry out self-adaptive adjustment on control gain.
Compared with the prior art, the invention has the advantages that:
(1) The novel fixed time convergence disturbance compensation observer designed by the invention can ensure that the observation error is rapidly converged to zero in the fixed time, has good noise suppression capability, and greatly improves the anti-interference capability of the aircraft;
(2) The double-layer self-adaptive gain strategy provided by the invention can carry out self-adaptive adjustment on the control gain, so that on one hand, the control gain is as small as possible to reduce buffeting, and on the other hand, the control gain is large enough to ensure the rapid convergence of the sliding mode surface. In addition, the method does not require maximum and minimum allowable values of gain and boundary information of the disturbance and its derivative.
Drawings
FIG. 1 is a flow chart of the present invention.
Detailed Description
As shown in fig. 1, the invention provides a reentry vehicle attitude control method based on adaptive gain disturbance compensation, which comprises the following steps:
(1) Reentry kinematics and dynamics modeling
Reentry kinematics and dynamics of the reentry vehicle are established as follows:
wherein θ= [ α, β, μ ]] T α, β, μ represent angle of attack, sideslip angle, and roll angle, respectively; omega= [ omega ] x ,ω y ,ω z ] T ，ω x ,ω y ,ω z Representing three-channel attitude angular velocity; b represents a control matrix, u represents a control input vector; d represents disturbance moment, and ΔJ represents a moment of inertia error matrix; the matrix R, J, Ω expressions are:
wherein J is x ，J y ，J z Three-channel main moment of inertia, J xy ，J xz ，J yz Respectively the product of inertia between the two channels.
Further simplification of the kinetic model can be obtained:
wherein the method comprises the steps ofRepresenting the total disturbance term, assuming that the boundary is an unknown constant, i.e. delta D.
(2) And establishing a fixed time convergence disturbance compensation observer, and observing the state and disturbance item of the reentry vehicle.
The novel fixed time convergence observer is of the form:
wherein z is 1 ,z 2 The observed amounts of ω, Δ,z respectively 1 ,z 2 Is the derivative of the power exponent alpha i And beta i The construction mode of (a) is as follows: alpha is E (1-epsilon, 1) and alpha i ＝iα-(i-1)，i＝1,2，β∈(1,1+ε 1 ) And beta is i ＝iβ-(i-1)，ε>0 and epsilon 1 >0 are all positive small amounts; gain kappa i And k i The matrix is constructed to satisfy the helvetz condition, here:
setting an observer observation error e 1 ＝z 1 -ω，e 2 ＝z 2 - Δ, then the error equation of the fixed time converging disturbance compensation observer is as follows:
here, theE is 1 ,e 2 Is a derivative of (a).
The observer convergence theorem is given below: in an observation error system (5), error vectorsConvergence to origin in a fixed time, i.e. observer state quantity +.>Converging to [ omega ] at a fixed time T Δ T ] T And the convergence time satisfies:
wherein ρ=1- α 1 ，σ＝1-β 1 ，r＝λ min (Q)/λ max (P)，r 1 ＝λ min (Q 1 )/λ max (P 1 )，Υ≤λ min (P 1 ) Positive, Q and Q 1 Are all 2 x 2-dimensional symmetric positive definite matrices, matrices P and P 1 Satisfies the Lyapunov equation:
wherein A and A 1 Has been given by equation (5).
The observer convergence proof is given below:
consider first the high power part of error equation (5):
due to matrix A 1 Satisfying the Hurwitz condition, and therefore the systemIs asymptotically stable, V 1 ＝e T P 1 e is the Lyapunov function of the linear system. For V 1 Deriving the time while considering equation (7), one can obtain:
the right half of the system (9) is continuous with respect to beta, so when beta epsilon (1, 1+ epsilon) 1 ) And epsilon 1 >0 is sufficiently small, V 1 (ψ)＝ψ T P 1 Psi also satisfiesWherein->Thus V 1 (ψ) is the Lyapunov function of the system (9). It should be noted that the right side of the system (9) is Ji Cidu m 1 ＝β-1>0 and an expansion weight s i Vector field of = (i-1) β - (i-2), i=1,..n. Under the condition that beta is close enough to 1, lyapunov function V 1 (ψ) Ji Cidu is l 1 ＝1>max(-m 1 0), and->Ji Cidu is l 1 +m 1 ＝1+m 1 And have the same dilation weight. Thus:
wherein, c 1 Is positive, 1+m 1 >1。
For linear systemsThe Rayleigh inequality and equation (9) mean V 1 (e)≤λ max (P)||e|| 2 And->Thus, it is possible to obtain:
where δ is any small amount. Considering that equation (9) is continuous with respect to β, the following inequality is shown:
this is true.
Taking into account parameter y>0 and V 1 (ψ(t 0 ))>And y. Calculate V 1 (ψ) knowing V at the full derivative of the system (5) 1 Will be reduced to gamma and for no longer than. Thus there is:
further, consider a low power system:
definition of the definitionScalar function V (ζ) =ζ T And P. Consider the publicFormula (13), then at V 1 The value of V (ζ) at decrease to γ satisfies:
V(ξ)≤λ max (P)||ξ|| 2 ≤λ max (P) (15)
since matrix A satisfies the Helviz condition, a linear systemIs asymptotically stable. Thus, the function v=e constituted by Lyapunov matrix P T Pe is the system->Is a Lyapunov function of (a). The derivative of this function satisfies the inequality:
when alpha is within (1-epsilon, 1)>0 and small enough), for V (ζ) =ζ T The same holds true for pζ. Since the right side of the system (14) is the expansion weight r i = (i-1) α - (i-2), (i=1,., n), the homogeneity is a homogeneity vector with m=α -1. Note that the vector field homogeneity is l=1 for the function V (ζ) under conditions that satisfy α sufficiently close to 1>max(-m,0)；Ji Cidu is l+m=1+m in the vector field>0, and the expansion weight is r i (i=1,) n. The following inequality holds:
wherein c is a normal number, 1+m <1.
For linear systemAs can be seen from the Rayleigh inequality, V (e). Ltoreq.lambda. max (P)||e|| 2 And is also provided withThus:
where δ >0. Considering that the system (14) has continuity due to the parameter alpha, the inequality
As it is, the constant c in the formula (17) is c=λ min (Q)/λ max (P). The corresponding convergence time isSo far, theorem proves complete.
(3) And (3) designing a controller: and designing an integral sliding mode controller, introducing an observer disturbance observation item into the integral sliding mode controller, and adopting a double-layer self-adaptive gain strategy to carry out self-adaptive adjustment on control gain.
Definition of tracking error e θ ＝θ-θ c Wherein θ is c And (3) representing a desired attitude angle instruction, and combining the formula (2) to obtain an attitude tracking error equation:
virtual control quantity omega c The design is as follows:
ω c ＝R -1 (θ c -k 1 e θ ) (21)
where the gain k 1 For the parameters to be designed, an angular velocity tracking error e is defined ω ＝ω-ω c The angular velocity tracking error equation is:
the design controller is as follows:
u＝u 1 +u 2 (23)
wherein u is 1 For canceling known system constitution and external disturbances, u 2 For compensating the estimation error. Controller u 1 The design is as follows:
wherein B is + A pseudo-inverse matrix, z, representing a control matrix B 2 Expressing the observed quantity of the disturbance delta by the fixed time convergence observer, substituting the above formula (22) to further obtain an angular velocity error equation is:
wherein the method comprises the steps ofu 2 Consisting of nominal and adaptive parts, u is first 2 The nominal part design is as follows:
u 2n ＝-ρ n B + Je ω (26)
wherein ρ is n Is a variable to be designed, u 2 The specific form is as follows:
wherein the integrating slip-form surface s is designed as follows:
ρ (t) represents the outer layer adaptive control gain, r (t) represents the inner layer adaptive controlGain control and guarantee of a limited timeThe form is as follows:
wherein sign (·) is a sign function, r 0 Is the initial value of r (t), and gamma is the parameter to be designed and satisfies gamma>0, and delta (t) has the form:
wherein the method comprises the steps ofIs the filter output of equivalent control, alpha and epsilon are parameters to be designed and satisfy 0<α<1,ε>0。
(4) Stability demonstration
Selecting a Lyapunov equationThe derivative is as follows:
reselection of Lyapunov equationThe derivative is as follows:
from the above, it can be seen that V 2 Is bounded and thus can speculate that e θ ,s∈L ∞ Furthermore we can derive:
the integral of the two formulas can be obtained:
thus we can get e θ ∈L 2 ,s∈L 1 And due to s E L ∞ Thus can be inferred thatAs known from the Barbalat lements, e θ S is progressively stable.
The design of the reentry aircraft attitude control method based on the adaptive gain disturbance compensation is completed. The novel fixed time convergence disturbance compensation observer designed by the invention can ensure that the observation error is rapidly converged to zero in the fixed time, has good noise suppression capability, and greatly improves the anti-interference capability of the aircraft; meanwhile, the double-layer self-adaptive gain strategy provided by the invention can carry out self-adaptive adjustment on the control gain, so that on one hand, the control gain is as small as possible to reduce buffeting, and on the other hand, the control gain is large enough to ensure the rapid convergence of the sliding mode surface. In addition, the method does not require maximum and minimum allowable values of gain and boundary information of the disturbance and its derivative.
Claims (5)
1. A reentry aircraft attitude control method based on adaptive gain disturbance compensation is characterized by comprising the following steps:
step one: establishing a reentry flight kinematics and dynamics model of the reentry aircraft facing control;
step two: establishing a fixed time convergence disturbance compensation observer, and observing the state and disturbance item of the reentry vehicle;
step three: designing an integral sliding mode controller, introducing an observer disturbance observation item into the integral sliding mode controller, and adopting a double-layer self-adaptive gain strategy to carry out self-adaptive adjustment on control gain;
the reentry vehicle is a control-oriented reentry flight kinematics and dynamics model, specifically:
wherein θ= [ α, β, μ ]] T α, β, μ represent angle of attack, sideslip angle, and roll angle, respectively; omega= [ omega ] x ,ω y ,ω z ] T ，Is the first derivative; omega x ,ω y ,ω z Representing three-channel attitude angular velocity; b represents a control matrix, u represents a control input vector; d represents disturbance moment, and ΔJ represents a moment of inertia error matrix; the matrix R, J, Ω expressions are:
wherein J is x ，J y ，J z Three-channel main moment of inertia, J xy ，J xz ，J yz Respectively representing the inertia products between the two channels;
further simplifying reentry flight kinematics and dynamics models to obtain:
wherein the method comprises the steps ofRepresenting the total disturbance term, wherein the boundary of the total disturbance term is an unknown constant, namely the I delta I is less than or equal to D;
the fixed time convergence disturbance compensation observer is of the form:
wherein z is 1 ,z 2 The observed amounts of ω, Δ,z respectively 1 ,z 2 Is the derivative of the power exponent alpha i And beta i The construction mode of (a) is as follows: alpha is E (1-epsilon, 1) and alpha i ＝iα-(i-1)，i＝1,2，β∈(1,1+ε 1 ) And beta is i =iβ - (i-1), ε >0 and ε 1 All >0 are positive small amounts; gain kappa i And k i The matrix is constructed to satisfy the Helviz condition, where
2. The method for controlling the attitude of a reentry vehicle based on adaptive gain disturbance compensation according to claim 1, wherein: let the observation error be e 1 ＝z 1 -ω，e 2 ＝z 2 - Δ, then the error equation of the fixed time converging disturbance compensation observer is as follows:
here, theE is 1 ,e 2 Is a derivative of (a).
3. The reentry vehicle attitude control method based on adaptive gain disturbance compensation of claim 2, wherein: and step three: an integral sliding mode controller is designed, an observer disturbance observation item is introduced into the integral sliding mode controller, and meanwhile, a double-layer self-adaptive gain strategy is adopted to carry out self-adaptive adjustment on control gain, specifically:
(1) For reentry flight kinematics and dynamics models, virtual control quantity omega is designed c Is that
ω c ＝R -1 (θ c -k 1 e θ )
Where the gain k 1 For the parameters to be designed e θ ＝θ-θ c For tracking error, θ c Representing a desired attitude angle command;
(2) The integral sliding mode controller u is designed for reentry flight kinematics and dynamics models as follows:
u＝u 1 +u 2
wherein u is 1 For canceling known system constitution and external disturbances, u 2 For compensating the estimation error;
u 1 in the form ofB + A pseudo-inverse matrix representing the control matrix B;
u 2 in the form of
Wherein u is 2n ＝-ρ n B + Je ω To compensate for nominal part of the error, e ω Tracking error e for angular velocity ω ＝ω-ω c ，ρ n As a variable to be designed,for the integral slip plane, ρ (t) represents the outer layer adaptive control gain, and r (t) represents the inner layer adaptive control gain in the form:
wherein sign (·) is a sign function, r 0 For an initial value of r (t), γ is the parameter to be designed and satisfies γ >0, while δ (t) has the form:
wherein the method comprises the steps ofIs the filtering output of equivalent control, alpha and epsilon are parameters to be designed, and satisfy 0 < alpha <1 and epsilon >0.
4. A reentry vehicle attitude control system implemented in accordance with the adaptive gain disturbance compensation-based reentry vehicle attitude control method of claim 1, comprising:
a dynamics model modeling module: establishing a reentry flight kinematics and dynamics model of the reentry aircraft facing control;
the observer building module: establishing a fixed time convergence disturbance compensation observer, and observing the state and disturbance item of the reentry vehicle;
the controller design module: designing an integral sliding mode controller, introducing an observer disturbance observation item into the integral sliding mode controller, and adopting a double-layer self-adaptive gain strategy to carry out self-adaptive adjustment on control gain;
the reentry vehicle is a control-oriented reentry flight kinematics and dynamics model, specifically:
wherein θ= [ α, β, μ ]] T α, β, μ represent angle of attack, sideslip angle, and roll angle, respectively; omega= [ omega ] x ,ω y ,ω z ] T ，Is the first derivative; omega x ,ω y ,ω z Representing three-channel attitude angular velocity; b represents a control matrix, u represents a control input vector; d represents disturbance moment, and ΔJ represents a moment of inertia error matrix; the matrix R, J, Ω expressions are:
wherein J is x ，J y ，J z Three-channel main moment of inertia, J xy ，J xz ，J yz Respectively representing the inertia products between the two channels;
further simplifying reentry flight kinematics and dynamics models to obtain:
wherein the method comprises the steps ofRepresenting the total disturbance term, wherein the boundary of the total disturbance term is an unknown constant, namely the I delta I is less than or equal to D;
the fixed time convergence disturbance compensation observer is of the form:
wherein z is 1 ,z 2 The observed amounts of ω, Δ,z respectively 1 ,z 2 Is the derivative of the power exponent alpha i And beta i The construction mode of (a) is as follows: alpha is E (1-epsilon, 1) and alpha i ＝iα-(i-1)，i＝1,2，β∈(1,1+ε 1 ) And beta is i =iβ - (i-1), ε >0 and ε 1 All >0 are positive small amounts; gain kappa i And k i The matrix is constructed to satisfy the Helviz condition, where
Let the observation error be e 1 ＝z 1 -ω，e 2 ＝z 2 - Δ, then the error equation of the fixed time converging disturbance compensation observer is as follows:
here, theE is 1 ,e 2 Is a derivative of (a).
5. The reentry aircraft attitude control system of claim 4, wherein: an integral sliding mode controller is designed, an observer disturbance observation item is introduced into the integral sliding mode controller, and meanwhile, a double-layer self-adaptive gain strategy is adopted to carry out self-adaptive adjustment on control gain, specifically:
(1) For reentry flight kinematics and dynamics models, virtual control quantity omega is designed c Is that
ω c ＝R -1 (θ c -k 1 e θ )
Where the gain k 1 For the parameters to be designed e θ ＝θ-θ c For tracking error, θ c Representing a desired attitude angle command;
(2) The integral sliding mode controller u is designed for reentry flight kinematics and dynamics models as follows:
u＝u 1 +u 2
wherein u is 1 For canceling known system constitution and external disturbances, u 2 For compensating the estimation error;
u 1 in the form ofB + A pseudo-inverse matrix representing the control matrix B;
u 2 in the form of
Wherein u is 2n ＝-ρ n B + Je ω To compensate for nominal part of the error, e ω Tracking error e for angular velocity ω ＝ω-ω c ，ρ n As a variable to be designed,for the integral slip plane, ρ (t) represents the outer layer adaptive control gain, and r (t) represents the inner layer adaptive control gain in the form:
wherein sign (·) is a sign function, r 0 For an initial value of r (t), γ is the parameter to be designed and satisfies γ >0, while δ (t) has the form:
wherein the method comprises the steps ofIs the filtering output of equivalent control, alpha and epsilon are parameters to be designed, and satisfy 0 < alpha <1 and epsilon >0.
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Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US9296474B1 (en) * | 2012-08-06 | 2016-03-29 | The United States of America as represented by the Administrator of the National Aeronautics & Space Administration (NASA) | Control systems with normalized and covariance adaptation by optimal control modification |
| CN107918400A (en) * | 2017-11-14 | 2018-04-17 | 中国运载火箭技术研究院 | A kind of in-orbit operation relative position and attitude combination control method of re-entry space vehicle |
| CN109144084A (en) * | 2018-07-11 | 2019-01-04 | 哈尔滨工业大学 | A kind of VTOL Reusable Launch Vehicles Attitude tracking control method based on set time Convergence monitoring device |
| CN109542112A (en) * | 2019-01-08 | 2019-03-29 | 哈尔滨工业大学 | A kind of set time convergence disturbance rejection control method returning to flight for the reusable rocket of VTOL |
| CN110733668A (en) * | 2019-10-16 | 2020-01-31 | 中山大学 | active vibration suppression method for flexible satellite based on component force synthesis and fixed time convergence |
| CN110844122A (en) * | 2019-11-20 | 2020-02-28 | 中山大学 | Satellite attitude control method for restraining attitude tracking error |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US9625913B2 (en) * | 2014-12-09 | 2017-04-18 | Embry-Riddle Aeronautical University, Inc. | System and method for robust nonlinear regulation control of unmanned aerial vehicles synthetic jet actuators |
-
2020
- 2020-11-20 CN CN202011314570.XA patent/CN112363524B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US9296474B1 (en) * | 2012-08-06 | 2016-03-29 | The United States of America as represented by the Administrator of the National Aeronautics & Space Administration (NASA) | Control systems with normalized and covariance adaptation by optimal control modification |
| CN107918400A (en) * | 2017-11-14 | 2018-04-17 | 中国运载火箭技术研究院 | A kind of in-orbit operation relative position and attitude combination control method of re-entry space vehicle |
| CN109144084A (en) * | 2018-07-11 | 2019-01-04 | 哈尔滨工业大学 | A kind of VTOL Reusable Launch Vehicles Attitude tracking control method based on set time Convergence monitoring device |
| CN109542112A (en) * | 2019-01-08 | 2019-03-29 | 哈尔滨工业大学 | A kind of set time convergence disturbance rejection control method returning to flight for the reusable rocket of VTOL |
| CN110733668A (en) * | 2019-10-16 | 2020-01-31 | 中山大学 | active vibration suppression method for flexible satellite based on component force synthesis and fixed time convergence |
| CN110844122A (en) * | 2019-11-20 | 2020-02-28 | 中山大学 | Satellite attitude control method for restraining attitude tracking error |
Non-Patent Citations (7)
| Title |
|---|
| Disturbance observer based fixed time sliding mode control for spacecraft proximity operations with coupled dynamics;Sixi Wu;《Advances in Space Research》;全文 * |
| Fixed-time disturbance observer based fixed-time back-stepping control for an air-breathing hypersonic vehicle;Xiao Wang;《ISA Transactions》;全文 * |
| Fixed-Time Disturbance Observer Based Nonsingular Fast Terminal Sliding Mode Guidance with Impact Angle Constraint;Rong Wu;《2018 IEEE CSAA Guidance, Navigation and Control Conference (CGNCC)》;全文 * |
| 垂直起降重复使用运载器返回制导与控制;韦常柱;《航空学报》;全文 * |
| 火箭垂直返回双幂次固定时间收敛滑模控制方法;崔乃刚;《哈尔滨工业大学学报》;全文 * |
| 重复使用运载器再入动力学建模研究;孙光;《导弹与航天运载技术》;全文 * |
| 高超声速飞行器再入自适应容错制导控制一体化设计;顾攀飞;《南京航空航天大学学报》;全文 * |
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